1. Field of the Invention
The present invention relates to reentry vehicles (RV). More specifically, the present invention relates to the field of low ballistic coefficient (BC) reentry vehicles (RV)—devices and systems that can be used to return objects and humans from orbit (e.g., low Earth orbit or LEO) to ground—and for aerobraking and aerocapture maneuvers in which a space vehicle, for example, in interplanetary space or in cis-lunar space in the Earth-Moon system, makes single or multiple passes through a planetary upper atmosphere to slow its velocity and allow it to enter into a desired orbit around that planet.
2. Description of the Related Art
The aerothermodynamic environment a reentry vehicle will experience is directly related to the velocity it is traveling at and the immediate density of the atmosphere along its reentry trajectory, i.e., where it is passing through. The altitude at which the predominant portion of deceleration takes place determines the density of the free stream atmosphere encountered by the vehicle. In general, the higher in the atmosphere that a reentry vehicle can commence deceleration, the lower will be the heat loading rate (expressed in watts/cm2).
Neglecting lift for the moment, the center of the entry corridor is defined by the nominal ballistic entry and determined by the vehicle's entry velocity and ballistic coefficient. The ballistic coefficient CB is a useful metric for assessing likely thermal loading during reentry and is expressed as the following equation:CB=(m/(Cd·A))  (I)where CB is the ballistic coefficient, m is the total mass of the vehicle, Cd is the vehicle drag coefficient, and A is the reference area, typically defined by maximum diameter.
For a given initial entry velocity, as the ballistic coefficient is increased, the vehicle descends deeper into the atmosphere and decelerates at lower altitudes. Thus, the ballistic coefficient is the dominant factor in determining the density profile that the vehicle encounters and significantly affects the severity of the aerothermodynamics environment.
The heating experienced by a vehicle is due to two modes of energy transfer from the flow field to the vehicle surface: convection and radiation. In general, convective heating is dominant for vehicles with a small nose radius, inversely proportional to the square root of the nose radius and directly proportional to the square root of the freestream density. Whereas, radiative heating is dominant for blunt vehicles at higher velocities where the flow field energy is sufficiently high, directly proportional to the nose radius and proportional to the freestream density to a power greater than one. Thus, for a given vehicle mass and entry velocity, heating can be reduced by decreasing the ballistic coefficient via increasing the vehicle size for a given mass. By reducing the ballistic coefficient, the vehicle flies a trajectory where a significant portion of the delta-V occurs higher in the atmosphere and, thus, at a lower density.
Convective heating is reduced by both the increased nose radius resulting from the larger vehicle size and the decrease in density. Radiative heating is potentially increased by the increase in nose radius but this effect is more than offset by the decrease in density resulting from the higher altitude trajectory. Therefore, to reduce the total heating to which a vehicle is subjected, it is desirable have an entry configuration with a low ballistic coefficient. The added benefit of a low ballistic coefficient entry configuration is that such configuration can reduce the overall heating to the point where a reusable thermal protection system (TPS) is achievable. Reduction of overall heating (to where TPS is achieved) is one of the primary objectives of the present invention. Importantly, a crucial and novel addition is the ability to dynamically maneuver throughout the reentry trajectory.
None of the vehicles in the prior art are dynamically maneuvering, that is, under active control, during reentry. They instead follow a ballistic trajectory defined by, among other things, the variances of the temperature and pressure of the atmosphere, the entry angle and ballistic coefficient. As a result, the landing footprint for such vehicles tends to be large. It is an explicit objective of the proposed invention to reduce the circular error probable of the landing zone to the smallest achievable number, nominally to within several hundred meters of a designated landing zone.
There are two general classes of designs in use today for reentry thermal protection systems (TPS): rigid systems and non-rigid systems. The vast majority of vehicles designed to date have relied upon rigid TPS. For small payloads it is possible to use a heat sink design. This approach, which utilizes thick segments of refractory metals for the TPS, has been in use for decades for data recovery from reconnaissance satellites requiring return of film canisters. This design employs a preponderant mass within the absorber shell and is considered untenable for large payload return to Earth. Ablative TPS were used for all United States manned missions through Apollo and have been used for unmanned and manned missions, such as the Galileo probe to Jupiter and the Space Exploration Initiative (SEI) probe to Mars, respectively, where reusability was not a design criterion.
The second general class of TPS is a nonrigid/deployable system. These systems offer the advantage of being designable with a low enough ballistic coefficient to reap the benefits stated previously. A subset of this class is inflatable systems. The concept of an inflatable aerodynamic decelerator has been around since the 1950's and variations of these so-called “ballutes” (balloon-parachutes) have been used in aviation and have been tested in high altitude sounding rockets.
Inflatable concepts have been exploited in a number of recent designs, the most publicized of which were the 2000 and 2002 launches of the Russian-Daimler-Benz Aerospace (DASA) and Russian-European Space Agency (ESA) Inflatable Reentry and Descent Technology (IRDT)-1 and IRDT-2 micro RV secondary payloads. The IRDT aeroshells measured 800 mm diameter and 700 mm height in the stowed configuration and 3.8 m diameter in the fully inflated configuration. The Russian Fregat primary upper stage in the 2000 test was itself returned via an inflatable aeroshell. The reentry first stage inflation achieved a diameter of 8 m with a second stage decelerator expanding to 14 m.
The first IRDT mission was performed on Feb. 9, 2000, by using the first Soyuz-Fregat test flight as a piggy-back launch opportunity. The results of this experiment were mixed. After launch from Baikonur and six orbits, the re-entry sequence was initiated and the IRDT vehicle apparently successfully deployed the first stage inflatable TPS but failed to deploy the second stage and crashed at >60 m/s. The Fregat system, initially lost, was recovered but had been vandalized and only limited information was obtained.
IRDT-2 was launched on a suborbital flight on Jul. 12, 2002 from a Russian submarine in the Barents Sea near Murmansk on board a converted Volna SS-N-18 intercontinental ballistic missile. However, due to a failure in the launcher/payload interface, IRDT-2 did not land in the expected nominal area on the Kamchatka peninsula and could not be activated. Further conclusions towards the IRDT system were not possible.
The objective of the second IRDT test flight was to verify the enhanced system concept under representative orbital conditions (7 km/s entry velocity, −6.9 degrees entry angle). In accordance with the recommendations of the IRDT-1 flight evaluation, certain refinements were implemented, e.g., improvements to the shield design and the internal pressure control and monitoring, introduction of a telemetry system for the landing phase and an enriched sensor package providing the flight evaluation data. The total mass amounted to approx. 140 kg.
Although the performance of this system remains to be firmly established, largely due to flight deployment problems, follow-on variations have been proposed by the Russian-ESA alliance for return of payloads from the ISS.
Similar designs were prototyped by NASA Langley over the past decade under the name Inflatable Reentry Vehicle Experiment (IRVE) and flown on Black Brant sounding rocket sub-orbital tests with ultimately successful results with temperatures experienced during the flights closely following theory and validating the predicted aerothermal behavior of low ballistic coefficient reentry vehicles.
In the United States, inflatable ballutes have been proposed for several planetary missions. For example, the Mars Surveyor (2003/2005) ballute study, conducted by ILC Dover and LMA), has considered the design of a toroidal inflatable ballute and attached aeroshell that is jettisoned once the vehicle achieves the desired velocity shedding. The ballute was designed for a 300-second operation duration, a maximum dynamic pressure 1.18 psi and an internal pressure of 2.4 psi. Zylon® 2000 was selected as the primary fabric for the ballute and was coated with LT-50 silicone. The TPS consisted of two outer layers of Nextel® 312 and 1 inner layer of Fiberfrax ceramic felt.
Although inflatable systems can be a viable option for a single-use aerocapture or entry, descent and landing TPS system, reusability is questionable. The major concern is the risk associated with the ability to stow the system between usages to protect the inflatable aeroshell from debris and micrometeorite damage.
Another option for a deployable non-rigid TPS system is the reentry aeroshells developed by ILC Dover and Ball Aerospace for candidate Mars micro-mission scenarios. Inflatable stiffeners were used to deploy a high temperature fabric. A second approach was developed using sprung composite rods to “unfurl” the stowed aerobrake. A final example of deployable reentry vehicles using stretched fabric is the canister deployment of the second stage of the NASA SOAREX reentry vehicle known as a Tube Deployed Reentry Vehicle (TDRV). The TDRV uses four stiffening panels that unfold when the vehicle is ejected from its carrying canister. These stiffening panels in turn serve as semi-rigid restraints for four lobes of a refractory fabric that trail behind the vehicle. Unlike the present invention, the prior art vehicles just described have the bulk of the mass of the reentry vehicle in front of the deployed drag surface.
To make orbital operations (e.g., in low Earth orbit—LEO) safer and more commercially viable for the private sector, a means of bringing mass back down to Earth more frequently and at much lower cost is required. Such a capability would be enabling for low- and zero-gravity manufactured products, e.g., pharmaceuticals, requiring regular delivery to customers who cannot wait the lengthy periods that presently exist with government-sponsored space program vehicles. Accordingly, there is a need for an operational on-demand, low cost, compact, small landing footprint capability will provide nascent space-manufacturing and space tourism industries with an off-shelf option for on-demand return of products, and for improving mission survivability and the security of their paying passengers.